Sun sensors are widely used in spacecraft attitude determination subsystems to provide a measurement of a sun vector in a spacecraft coordinate. They have wide applications in attitude measurement and control of satellite and other space vehicles. A novel complementary metal oxide semiconductor (“CMOS”) sun sensor includes: an optical system with single aperture or multi aperture, electronic system and process computer based on a field programmable gate array (“FPGA”) or an advanced RISC machines (“ARM”). The principle of sun sensors is that the sunlight is incident on the CMOS image sensor, and sunspots are formed through the apertures of an optical system. The positions of sunspots vary with different incidence angles. Then image processing and attitude calculation are done by an electronic system and process computer to provide the corresponding attitude angle of the satellite.
The operation mode of the developed CMOS sun sensor is: image grabbing, simple image processing and communication with process computer, which is accomplished by the electronic system; further image processing (e.g. the centroid calculation, sunspot identification and attitude calculation) that is implemented by the process computer. The function of a kind of sun sensor electronic systems is: image grabbing, image segmentation and transmission of sunspot information above the threshold to the process computer. The function of another kind of electronic systems is only image grabbing while image processing and attitude calculating are accomplished by the process computer. In the two operation modes the burden of process computer is increased and the amount of image data to be transmitted to process computer is large because of full frame data transmission which limits the attitude update rate of sun sensor.
And moreover, when the aperture of the optical system is polluted and some apertures can not pass the sunlight, the attitude must still be calculated accurately, and the reliability of the sun sensor must be improved.